"After successfully completing the static test programme on the F-35 (known as AG-1) we have now returned the aircraft back to Lockheed Martin, Fort Worth.

The static test programme broke all records for the speed of testing having applied more than 150 different loading configurations in just over nine months.

Having proven the strength of the aircraft is now beginning the 4500 mile journey back to the US after almost three and a half years in the structural test facility at Brough.

Static testing the F-35 means that the aircraft has been ‘flown’ to its limits with loads applied to it replicating the effect of high gravitational forces far beyond any conditions likely to be flown in actual flight. This is done with the airframe nesting in a multi-million pound rig kitted out with over 4000 strain gauges, 170 actuators and over 50 miles of wiring at our Brough site in Yorkshire. Brough is home to a world leading facility for putting aircraft through their paces to ensure they are strong enough and resilient enough to perform in the harshest environments in the world.

Tim Bramhall runs the F-35 structural test programme at Brough said “We certainly don’t give the aircraft an easy ride here. We push it to its limits so that we can be confident that each of the 3000+ aircraft that have been ordered will perform safely and effectively. The real challenge is keeping aircraft weight at a minimum whilst maintaining the strength of the plane within certain specified limits”

With this set of tests complete Tim added “We still have another F-35 CTOL airframe in the facility undergoing fatigue testing along with the remaining horizontal and vertical tails from the Carrier variant. Work on those continues on schedule and are shining examples of the long term future the structural test facility has ahead.”

All very nice, I'm sure, but there is a big hole in the F-35 static test logic. It is important to understand the differences between design loads, static test loads, and flight test loads. Design loads, which are used to design the airplane structure, are based on preliminary weight distribution, air pressure distribution, flight control laws, and flexibility effects. Years later, the static test loads are derived from updated values from all those items and may be quite different from design loads. Flight test measured loads are the real deal, as they eliminate all estimates and guesswork from analysis.

The usual way to conduct static test and flight test is to static test at 100% of updated design loads, then flight test at a conservative 80% limit. These 80% flight loads are extrapolated to 100% conditions and compared with 100% static test loads. If any extrapolated flight test load conditions are greater than 100% static test, the static test conditions are revised upward. Almost inevitably, that condition will occur. Then 150% static test conditions are applied to allow conservative 100% flight test conditions to be flown. Following a complete analysis of static test and flight test results, the fleet airplanes may be released to fly 100% conditions.

So where did F-35 static test logic go wrong? All static tests were completed to 150% before any flight test loads were available. So there will be no opportunity to revise static test loads upward if flight test loads show the need to do so. And by dismantling the test article updated conditions can never be applied.

If they are very lucky, flight test loads will be no larger than static test loads and all will be OK. If not, what will they do? I am really surprised the services, especially the Navy, allowed LM and BAE to follow that static test scheduling.

Hey John, I was just thinking that with the digital flight control system the computers will always ensure that the aircraft is flown with in its prescribed structural limits in order to prevent the pilot from bending or breaking the airplane.

Well I might point out that nowhere in that article does it mention anything about 150% of anything. That being said, the aircraft are stressed far past the DESIGN limits of the airframe until failures are detected...much like commercial aircraft. Johnwill's argument is invalid.

True enough, but flight control computers limit g or roll rate or control surface deflection, for example, but do not limit structural load. One purpose of the static test program is to demonstrate the strength of the entire structure is at least 150% of the largest load it will ever experience in flight. The problem is the static test conditions were based on analysis only. Analysis methods are certainly being improved all the time, but if they were perfect, there would be no reason to verify them with measured loads in flight test. There is a good chance flight test will show loads higher than 100% static test in some maneuver, at some mach/altitude, on some component of the airplane. So, you end up with a structure which has not been adequately static tested. What can be done? One way is to use more stress analysis to show an adequate strength margin, but that is somewhat risky. You can use the flight control computer to reduce maneuver limits, but that penalizes the performance of the airplane to less than spec values.

To clarify, here is a simple example. Let's say under some condition, flight test shows wing reaches 100% limit bending moment at 8.8g rather than the required 9g. Stress analysis might be able to show there is enough margin to cover the shortfall in the static test load. Or, the flight control computer could be programmed to limit the load factor to 8.8 under specific conditions.

Another problem is the effect of higher than planned loads on the long term durability of the structure, called fatigue in years gone by.

You might wonder why there is a 50% factor of safety required. It is to cover sub-standard materials, manufacturing tolerances and errors, cracks and delaminations which develop during thousands of flight hours, overload due to tolerances in the flight control computer sensors and controls, and inevitable weight growth over the years.

More than likely, the people who planned this program are a lot smarter than I am and have somehow met the requirements.

Last edited by johnwill on 13 Aug 2012, 03:48, edited 1 time in total.

As long as flight test loads are not 50% greater than predicted loads then there shouldnt be a concern. For stuff like door loads (e.g. AAI door issue) the structure can be locally updated and retested. At the end of the day that extra 50% margin is there for a reason; to account for inherently imperfect engineering predictions.

“Its not the critic who counts..The credit belongs to the man who does actually strive to do the deeds..”

checksixx wrote:Well I might point out that nowhere in that article does it mention anything about 150% of anything. That being said, the aircraft are stressed far past the DESIGN limits of the airframe until failures are detected...much like commercial aircraft. Johnwill's argument is invalid.

And the requirement is to demonstrate 150% of design limit in Static Test with NO failure. Please explain why you think my argument is invalid.

LMAggie wrote:As long as flight test loads are not 50% greater than predicted loads then there shouldnt be a concern. For stuff like door loads (e.g. AAI door issue) the structure can be locally updated and retested. At the end of the day that extra 50% margin is there for a reason; to account for inherently imperfect engineering predictions.

Sorry, your statement about no concern so long as flight test loads are not more than 50% greater than predicted is entirely incorrect.

We're not talking about secondary structure like doors, but primary structure, things like wings, tails, etc. I assure you any exceedance of predicted loads is cause for concern. The requirement remains, Static Test to 150% of verified loads without failure is required.

Inherently imperfect engineering predictions are the justification for flight loads test to obtain the "real" loads. I already explained the reasons for the 50% over-load capability, and they do not include imperfect engineering predictions.

LMAggie wrote:As long as flight test loads are not 50% greater than predicted loads then there shouldnt be a concern. For stuff like door loads (e.g. AAI door issue) the structure can be locally updated and retested. At the end of the day that extra 50% margin is there for a reason; to account for inherently imperfect engineering predictions.

Sorry, your statement about no concern so long as flight test loads are not more than 50% greater than predicted is entirely incorrect.

We're not talking about secondary structure like doors, but primary structure, things like wings, tails, etc. I assure you any exceedance of predicted loads is cause for concern. The requirement remains, Static Test to 150% of verified loads without failure is required.

Inherently imperfect engineering predictions are the justification for flight loads test to obtain the "real" loads. I already explained the reasons for the 50% over-load capability, and they do not include imperfect engineering predictions.

Based solely on Johnwill's credentials, I am inclined to believe his concerns. Every new material used in aircraft manufacturing presents new challenges for testing. An old example of this was the F-111: http://www.f-111.net/cold.htm

I hope the F-35 passes all the tests AND never encounters a failure.
fisk

I don't disagree about what the requirement is. My point is that the 50% is a margin for the unknown and just because the tested loads aren't 100% accurate, doesn't mean the aircraft isn't airworthy. If it did, then obviously they would have repeated legacy test practices.

“Its not the critic who counts..The credit belongs to the man who does actually strive to do the deeds..”

Right, it doesn't mean the airplane isn't airworthy, but it allows for the possibility that it is not. Here again are the reasons for the 50% factor :

You might wonder why there is a 50% factor of safety required. It is to cover sub-standard materials, manufacturing tolerances and errors, cracks and delaminations which develop during thousands of flight hours, overload due to tolerances in the flight control computer sensors and controls, and inevitable weight growth over the years.

Please note that load analysis errors are not included. If weight growth is significant, then another static test might be necessary. For example, the original F-16A/B static test was conducted in 1977. By the time the Block 30 airplanes were being produced it was necessary to conduct another complete static test. In that test, around 1985, there was a complete wing failure at 137% of design limit load. Block 30 airplanes were beefed up to cover that deficiency and Block 40 airplanes were totally redesigned.

Fiskerwad's link to the F-111 cold proof test gives a perfect example of material deficiencies. Two crewmen paid for that deficiency with their lives and many million dollars were spent to correct it.

I am sure the quality of the static test, including the test rig, the load system, data recording, load distribution, etc are all first rate. The only problem is using un-verified loads. Perhaps they covered that possibility by arbitrarily increasing the loads by a small percentage, say 5%. That should help cover load prediction errors, yet allow the early completion of static test for whatever reason they want, probably cost.